Method of manufacturing turbine airfoil and tip component thereof using ceramic core with witness feature

ABSTRACT

Methods of manufacturing or repairing a turbine blade or vane are described. The airfoil portions of these turbine components are typically manufactured by casting in a ceramic mold, and a surface made up of the cast airfoil and at the least the ceramic core serves as a build surface for a subsequent process of additively manufacturing the tip portions. The build surface is created by removing a top portion of the airfoil and the core, or by placing an ultra-thin shim on top of the airfoil and the core. The overhang projected by the shim is subsequently removed. These methods are not limited to turbine engine applications, but can be applied to any metallic object that can benefit from casting and additive manufacturing processes. The present disclosure also relates to finished and intermediate products prepared by these methods.

The present disclosure generally relates to a method of manufacturing orrepairing a hollow metal object. More specifically, the hollow metalobject is prepared using an additive manufacturing (AM) technique, or amixture of an AM technique and of an investment casting technique. TheAM technique utilized in the manufacturing method is not limited to thedirect metal laser melting (DMLM) or any other laser powder-bed fusionadditive manufacturing. The hollow metal object produced is especiallyuseful as a component of an aircraft engine or other power generationturbines, e.g. a blade or a stator vane.

BACKGROUND

Superalloy materials are among the most difficult materials to weld dueto their susceptibility to weld solidification cracking and strain agecracking. The term “superalloy” as used herein means a highly corrosionand oxidation resistant alloy with excellent mechanical strength andresistance to creep at high temperatures. Superalloys typically includehigh nickel or cobalt content. Examples of superalloys include alloyssold under the trademarks and brand names Hastelloy, Inconel alloys(e.g., IN 738, IN 792, IN 939), Rene alloys (e.g., Rene N5, Rene 80,Rene 142, Rene 195), Haynes alloys, Mar M, CM 247, CM 247 LC, C263, 718,X-750, ECY 768, 282, X45, PWA 1483 and CMSX (e.g. CMSX-4) single crystalalloys.

Gas turbine airfoils, both rotating blades and stationary vanes, areoften manufactured by casting a superalloy material around a fugitiveceramic core that is then removed to form cooling chambers and channelsin the blade. The manufacture of these turbine blades, typically fromhigh strength, superalloy metal materials, involves an initialmanufacture of a precision ceramic core to conform to the intricatecooling passages desired inside the turbine blade. A precision die ormold is also created which defines the precise 3-D external surface ofthe turbine blade including its airfoil, platform, and integraldovetail. A schematic view of such a mold structure 10 is shown inFIG. 1. The ceramic core 11 is assembled inside two die halves whichform a space or void therebetween that defines the resulting metalportions of the blade. Wax is injected into the assembled dies to fillthe void and surround the ceramic core encapsulated therein. The two diehalves are split apart and removed from the molded wax. The molded waxhas the precise configuration of the desired blade and is then coatedwith a ceramic material to form a surrounding ceramic shell 12. Then,the wax is melted and removed from the shell 12 leaving a correspondingvoid or space 13 between the ceramic shell 12 and the internal ceramiccore 11 and tip plenum 14. Molten superalloy metal is then poured intothe shell to fill the void therein and again encapsulate the ceramiccore 11 and tip plenum 14 contained in the shell 12. The molten metal iscooled and solidifies, and then the external shell 12 and internal core11 and tip plenum 14 are suitably removed leaving behind the desiredmetallic turbine blade in which the internal cooling passages are found.In order to provide a pathway for removing ceramic core material via aleaching process, a ball chute 15 and tip pins 16 are provided, whichupon leaching form a ball chute and tip holes within the turbine bladethat must subsequently be brazed shut.

U.S. Patent Application Publication No. 2010/0200189 (assigned toGeneral Electric Company) discloses a method by which, as shown FIGS. 2Aand 2B, the outer end of the airfoil 18 may be closed. In a first step,as shown in FIG. 2A, a tip plate 50 that precisely defines that shape ofthe cross section of the airfoil 18 is placed on the outer end of theairfoil 18 (without the ceramic casting mold), in contact with the outerwall 19. The tip plate 50 is bonded to the outer wall 19 by laserwelding. Laser energy is then directed at the tip plate 50 from the endor the peripheral edges (see arrows “W” in FIG. 2A) so as to produce athrough-weld and fuse the outer periphery of the tip plate 50 to theouter wall 19. Next, as shown in FIG. 2B, the tip wall 34 is formedthrough a freeform laser fabrication process where molten alloy powderis deposited on the tip plate 50 in one or more passes.

In another aspect, U.S. Patent Application Publication No. 2010/0200189depicts an airfoil 18″ formed by an alternative method. FIG. 3Aillustrates the airfoil 18″ in the as-cast condition (without theceramic casting mold) with an outer wall 19″. The interior of theairfoil 18″ is filled with a suitable metallic alloy powder 68, which isscraped flush or otherwise leveled with the outer end of the outer wall19″. The powder 68 is sintered together and bonded to the outer wall 19″by directing laser energy at it, shown schematically at arrow “L” inFIG. 3A. FIG. 3B depicts the airfoil 18″ after the powder 68 has beensintered into a completed tip cap 36″ and the excess powder 68 removed.Once the tip cap 36″ is formed, a tip wall 34″ is formed on top of thetip cap 36″ using a freeform laser fabrication process, as shown in FIG.3C.

U.S. Patent Application Publication 2015/0034266A1 (assigned to SiemensEnergy, Inc.) describes a method of manufacturing a turbine blade wherethe cavity of the blade is also filled with a support material forsubsequent formation of the blade tip. As shown in FIG. 4, method 80includes step 82 where a superalloy turbine blade is initially castwithout a blade tip cap but with tip walls. At step 84, a supportingelement is placed in a cavity of the blade. Then at step 86, an additivefiller material comprising a metal powder is supported on the supportingelement across the opening. Next at step 88, an energy beam is traversedacross the additive filler material to melt the material and to therebyform a superalloy cap across the blade tip and is fused to the existingblade tip walls. The method 80 further comprises a step 90 of building aradially extending squealer ridge around the periphery of the cap viaadditive welding.

In view of the foregoing, and the fact that current blades and vanestend to be life limited especially at their tips which are veryexpensive to replace, a need remains for novel methods of manufacturingtips or other components for new-make cast airfoils and field-returnrepair airfoils. It would be desirable to provide methods that are lesstime-consuming and more cost-effective. For example, it would beespecially beneficial to provide methods that utilize materials that arealready existing in airfoil manufacturing facilities and/or methods thatfind new, secondary uses for such materials, thereby circumventing theneed to acquire and waste any new materials.

SUMMARY

In a first aspect, the present invention relates to a method ofmanufacturing a metal object. The method comprises: (a) pouring a liquidmetal into a ceramic casting mold to form a cast component uponsolidification of the liquid metal, wherein the ceramic casting moldcomprises a ceramic core that comprises a witness feature and whereinthe ceramic core fills an inner cavity of the cast component; (b)removing a portion of the ceramic casting mold and the cast componentuntil the witness feature is revealed to create a surface portion ofcast component and a surface portion of the ceramic core; (c) depositinga layer of metallic powder onto the surface portions; (d) irradiating atleast a portion of the metallic powder to form a fused layer; and (e)repeating steps (c)-(d) until the blade or vane is formed. Preferably,the surface portions are planar surfaces.

In one embodiment, the ceramic casting mold further comprises a ceramicshell and at least one inner cavity between the shell and the core.Accordingly, an embodiment of the method further comprises removing theceramic shell before step (b).

In certain embodiments, the method further comprises treating thesurface portions to prevent ceramic contamination.

In one embodiment, steps (c)-(e) of the method are carried out in thepresence of induction heating, radiant heating or a combination of both.

In some embodiments, the method further comprises removing the ceramiccasting mold from the metal object.

Preferably, the metal object is a turbine component. Preferably, theturbine component is a blade or a vane, and the cast component is anairfoil.

In a second aspect, the present invention relates to method ofmanufacturing a turbine blade or vane.

In a third aspect, the present invention relates to cast metal turbineblade or vane workpiece comprising a metallic blade or vane portion witha ceramic core that comprises a witness feature; and a substantiallyplanar top surface of the workpiece that comprises a metallic portionand a ceramic portion.

In one embodiment, the ceramic core fills an inner cavity of themetallic blade or vane.

In one embodiment, the blade or vane is further surrounded by a ceramicshell.

In one embodiment, the cast metal turbine blade further comprises aprotective layer for preventing ceramic contamination.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic diagram showing an example of a conventionalscheme for a core-shell mold with ball chute prepared by a conventionalinvestment casting process.

FIGS. 2A and 2B are schematic diagrams showing a prior art method ofmanufacturing a blade tip where a tip plate is added to an as-castairfoil (ceramic casting mold removed) the tip wall is built upon thetip plate.

FIGS. 3A, 3B and 3C are schematic diagrams showing another prior artmethod of manufacturing a blade tip where an as-cast airfoil (ceramiccasting mold removed) has its hollow interior filled with a metallicalloy powder until the powder is leveled with the outerend of the outerwall, sintered to form a tip cap, then the tip wall is built upon thetip cap.

FIG. 4 is a flowchart depicting yet another prior art method of a bladetip where an as-cast airfoil (without the ceramic casting mold) has itshollow interior filled with a supporting element followed by an additivefiller material across the blade tip. An energy is applied across theadditive filler material to melt the material to form a superalloy capacross the blade tip and is fused to the existing blade tip walls.

FIG. 5A shows a ceramic casting mold having an external shell and aninternal core with one or more inner cavities between the shell and thecore according to an embodiment of the invention.

FIG. 5B shows the inner cavities of the ceramic casting mold of FIG. 5Abeing filled with a liquid metal to cast an airfoil.

FIG. 5C shows a portion of the liquid metal-filled ceramic casting moldof FIG. 5B being ground to create a flat or planar surface.

FIG. 5D depicts a DMLM process of manufacturing a blade tip according toan embodiment of the invention, where the apparatus is equipped with anexternal heat control mechanism.

FIG. 5E is an expanded view of the DMLM process of FIG. 5D showing themetal-ceramic intermediate being supported by the component translatingmechanism and further secured by seals, and with the external heatingmechanisms in proximate positions to the intermediate.

FIG. 5F depicts a DMLM process of manufacturing blade tip according toan embodiment of the invention, where the DMLM system is equipped with adirect ink write (DIW) lithography system for deposition of ceramicslurry.

FIG. 6A shows a ceramic casting mold having an external shell and aninternal core with one or more inner cavities between the shell and thecore according to an embodiment of the invention, where the core has anembedded witness feature.

FIG. 6B shows the ceramic casting mold of FIG. 6A (filled with liquidmetal) being ground until the witness feature is exposed.

FIG. 7A is a cutaway view of a turbine blade (with the aft side of theblade looking forward) having a plurality of turbulators orientedcircumferentially .

FIG. 7B is a cutaway view of a turbine blade (with the aft side of theblade looking forward) having a plurality of turbulators orientedaxially.

FIG. 8A presents a prior art ceramic casting mold having a tip plenumcore and tip pins.

FIG. 8B presents a ceramic casting mold having an open tip castingaccording to an embodiment of the present invention.

FIG. 9A shows a field-return turbine blade with a damaged tip.

FIG. 9B shows restoration of the internal ceramic core and optionallythe external shell by injection of a ceramic slurry into the hollowfield-return turbine blade of FIG. 9A.

FIG. 10 depicts a method of repairing a field-return turbine bladeaccording to an embodiment of the present invention.

DETAILED DESCRIPTION

The detailed description set forth below in connection with the appendeddrawings is intended as a description of various configurations and isnot intended to represent the only configurations in which the conceptsdescribed herein may be practiced. The detailed description includesspecific details for the purpose of providing a thorough understandingof various concepts. However, it will be apparent to those skilled inthe art that these concepts may be practiced without these specificdetails. For example, the present invention provides a preferred methodfor additively manufacturing certain components of metal objects, andpreferably these components and these objects are used in themanufacture of jet aircraft engines. Specifically, the production ofsingle crystal, nickel-based superalloy or elemental titanium hollowmetal objects such as turbine blades and stator vanes can beadvantageously produced in accordance with this invention. However,other metal components of the turbine may be prepared using thetechniques and methods described herein. Similarly, other suitable,non-turbine components may also be prepared using the techniques andmethods provided herein.

As established in the background, it is known that the tip portion of ablade or a stator vane can be manufactured, subsequent to casting of theairfoil and other portions (e.g. blade root and vane trunnions), by adirect metal laser melting (DMLM) or an electron beam melting (EBM)process such as the EBM processes described in U.S. Pat. No. 9,064,671(assigned to Arcam AB and incorporated herein by reference in itsentirety).

A description of a typical DMLM process is provided in German Patent No.DE 19649865, which is incorporated herein by reference in its entirety.Owing to the fact that DMLM process requires a build platform or surfacefor supporting fabrication of a desired part geometry, prior art methodshave often required building of a custom-made tip plate or tip cap (i.e.precisely defining that shape of the cross section of the airfoil) tonot only close off the hollow interior of the airfoil, but to also serveas the surface for the DMLM process.

Ceramic casting mold used in conventional investment casting processeshas often been regarded as a fugitive material to be removed immediatelyafter cooling and solidification of the poured liquid metal. The presentinventors have discovered that the internal ceramic core and optionallythe external shell can elegantly and conveniently serve as a supportstructure for subsequent DMLM formation of the tip or any othercomponent of the blade or vane. As used herein, the term “fugitive”means removable after melting and cooling of the metal, for example by amechanical process, by fluid flushing, by chemical leaching and/or byany other known process capable of removing the fugitive material fromits position. In certain embodiments when working with field-returnairfoils with damaged or worn-off tips, ceramic slurry can be injectedinto the hollow interior or inner cavity of the airfoils to re-form theceramic core. The combined core and airfoil then offers a supportsurface for formation of a replacement tip by DMLM.

The use of the ceramic core as a support structure for subsequentadditive manufacturing operations is not only cost- and time-effective,but also highly viable. This is because ceramic is chemically inert andhas high strength, high fracture toughness, high elastic modulushardness, high melting points, low thermal expansion, excellent wearresistance, etc. Such physicochemical properties make ceramic an idealmaterial to withstand the conditions (i.e. high temperature and highpressure) of additive manufacturing processes. Moreover, it is notdifficult to remove the ceramic core after completion of themanufacturing process (i.e. casting and printing), for example bymechanical force or chemical leaching (e.g. in alkaline bath) orpreferably a combination of both. The ceramic casting molds, cores,shells and slurries of the present invention are preferably composed ofa photopolymerized ceramic, more preferably a cured photopolymerizedceramic.

FIGS. 5A-5D depict a method of manufacturing a new-cast turbine blade inaccordance with the present invention. In the beginning, as shown inFIG. 5A, there is provided a ceramic casting mold 500 having an externalshell 502 and an internal core 504. The cavities 506 are adapted todefine the shape of the turbine blade upon casting and removal of theceramic casting mold 500. During the casting process, as seen in FIG.5B, a liquid metal 508 is poured in the cavities 506 and left to cooland solidify to form at least the airfoil component 510 of the turbineblade. In some embodiments, the root of the turbine blade or the innerand outer trunnions of a stator vane are also formed during casting (notshown). Next, as shown in FIG. 5C, a portion of the airfoil 510-ceramiccasting mold 500 combination is removed, for example, by grinding,machining, cutting or other known techniques. FIG. 5C shows anembodiment where a grinder blade or disc is used remove the top portionof the airfoil 510-ceramic casting mold 500 combination. Consequently, aflat or planar surface 512 is revealed and the metal-ceramicintermediate 516 is produced. The planar surface 512 functions as acontinuous and sealed support structure for the subsequent deposition ofraw metallic powder in the DMLM process of manufacturing the tipcomponent of the turbine blade shown in FIG. 5D.

The tip of a blade or a stator vane in accordance with the presentinvention includes a peripheral tip wall that is sometimes referred as asquealer tip and a tip cap that closes off the interior of the airfoil.In some embodiments, however, the tip cap may be eliminated leaving theairfoil with an open cavity.

At the initial stage of the DMLM process, the raw metallic powderdeposited on the ceramic-metal build surface 512 is melted together andbonded to the outer wall of the airfoil 510 by directing laser energy atit. The exact process parameters may vary to suit a specificapplication. In one embodiment, a short pulsed infrared laser beam isused, with an average power of 1-100 W, pulse frequency of 1 Hz to 200kHz. The translation speed or scanning speed, if the laser beam is usedwith a scanner, is approximately 5 mm/s (0.197 in./s) to about 500 mm/s(19.7 in./s).

Alternatively and preferably, the first additive layer of the tip isjoined to the airfoil 510 using a method of bonding superalloysdisclosed in U.S. Pat. No. 8,925,792 (assigned to General ElectricCompany), which is incorporated herein by reference in its entirety. Themethod generally includes aligning a first superalloy subcomponenthaving a gamma-prime solvus g′1 and a second superalloy subcomponenthaving a gamma-prime solvus g′2, with a filler material that includes atleast 1.5 wt % boron disposed between the first and second superalloysubcomponents; performing a first heat treatment at a temperature T1,where T1 is above the solidus of the filler material and below theliquidus of the filler material; and performing a second heat treatmentat a temperature T2, where T2 is greater than T1, and where T2 isgreater than or equal to the lower of g′1 and g′2.

With proper control of the DMLM process parameters, this process canproduce the same microstructure in the additively manufactured tip wall(e.g. directionally solidified or single crystal) as that of the airfoil510, if desired. For example, a continuous wave beam of about 300 W toabout 1000 W power may be used, with a traverse rate of about 0.25 cm/s(0.1 in./s) to about 0.76 cm/s (0.3 in./s) and preferably about 0.44cm/s (0.175 in./s) to about 0.51 cm/s (0.200 in./s). About 100-200passes result in a tip wall of a suitable height and a near-net shape,where each layer is 20-100 μm, preferably 20-50 μm, more preferably30-50 μm. In one embodiment, each additive layer is 30 μm. As usedherein, the term “near-net” refers to a structure that does requiresubstantial additional machining processes in order to arrive at afinished part. Once the DMLM process is finished, the tip wall may befurther formed by known processes of machining, grinding, coating, etc.

Representative examples of suitable materials for the raw metallicpowder used during additive manufacturing and the liquid metal usedduring casting include alloys that have been engineered to have goodoxidation resistance, known “superalloys” which have acceptable strengthat the elevated temperatures of operation in a gas turbine engine, e.g.Hastelloy, Inconel alloys (e.g., IN 738, IN 792, IN 939), Rene alloys(e.g., Rene N4, Rene N5, Rene 80, Rene 142, Rene 195), Haynes alloys,Mar M, CM 247, CM 247 LC, C263, 718, X-750, ECY 768, 282, X45, PWA 1483and CMSX (e.g. CMSX-4) single crystal alloys. The airfoils and the bladetips of the present invention may be formed with one or more selectedcrystalline microstructures, such as directionally solidified (“DS”) orsingle-crystal (“SX”). In certain embodiments, the blades and vanes ofthe present invention may be formed with multiple, controlled metallicgrain orientations, such as the microstructures disclosed in theApplicant's co-pending application Ser. No. 15/405,656, of which thedisclosure is incorporated herein by reference in its entirety. In aparticular embodiment, the airfoils of the present invention are formedwith a single crystal structure but the blade tips are formed with anon-single crystal structure.

FIG. 5D shows an embodiment of the DMLM process where the buildenclosure 551 includes induction heating at least a portion of the bladetip that is being additively manufactured. The additive manufacturingapparatus 550 includes a build enclosure 551 that encloses, at leastpartially, components of the apparatus 550. For example, at least apowder bed (not shown) is provided within the build enclosure 551 suchthat fusion of metallic powder P in the powder bed occurs in a definedenvironment. In some embodiments, the build enclosure 551 defines anatmosphere that is substantially oxygen-free. In some embodiments, thebuild enclosure 551 defines an inert atmosphere (e.g., an argonatmosphere). In further embodiments, the build enclosure 551 defines areducing atmosphere to minimize oxidation. In yet further embodiments,the build enclosure 551 defines a vacuum.

As shown in FIG. 5D, the build enclosure 551 includes a first air-tightzone 552 that defines the environment in which the fusion of metallicpowder P in the powder bed occurs. The build enclosure 551 may alsoinclude a second zone 553 that may or may not be air-tight, and in oneexample defines an environment that is in communication with the firstzone 552. The apparatus 550 also includes a powder supply mechanism 554,a powder translating mechanism 555, a build platform 556, a work surface562, a build plate 557, a seal 558, an external heat control mechanism559 and a component translating mechanism 560. During production, anelevator 561 in the powder supply mechanism 554 lifts a prescribed doseof metallic powder “P” above the level of the build platform 556. Theprescribed dose of metallic powder is then spread in a thin, even layerover the work surface 562 and the build surface 512 to form the powderbed 565, by the powder translating mechanism 555. Overflows from thework surface 562 are collected by the excess powder receptacle 563, thenoptionally treated to sieve out rough particles before re-use.

The build plate 557 includes an aperture 564 extending from the buildplate 557. The aperture 564 may be in communication with, and partiallyform the powder bed 565. As the metallic powder “P” is pushed across theaperture 564 of the build plate 557 by the powder translating mechanism555, the metallic powder “P” is able to fall through the aperture 564and into the powder bed 565. In this way, the powder translatingmechanism 555 may be operable to deposit the metallic powder “P” throughthe aperture 564 of the build plate 557 and into the powder bed 565.

Preferably, the build plate 557 is made of a substantiallynon-conductive material (e.g.

ceramic, glass or otherwise non-metallic) that is operable to preventthe external heat control mechanism 559 (which is operable to form apredetermined temperature profile of the intermediate 516) from heatingthe build plate 557 to a sintering temperature of the metallic powder“P” that is utilized to form layers of the blade tip added to theairfoil 510. In this way, the external heat control mechanism 559 isable to form a predetermined temperature profile of the airfoil 510 andthe to-be manufactured tip without interfering with the fusion orsintering of the metallic powder “P” that forms layers of the tip.Preferably, the external heat control mechanism 559 is induction-based,but radiant- or laser-based heating may also be used (e.g. with heatlamp(s) or auxiliary laser).

The external heat control mechanism 559 in the example shown in FIGS. 5Dand 5E is positioned proximate to the bottom surface 566 of the buildplate 557. In some embodiments, as shown in FIG. 5D, the external heatcontrol mechanism 559 may be spaced from the bottom surface 566 of thebuild plate 557. In an alternative embodiment, the external heat controlmechanism 559 may abut the bottom surface 566 of the build plate 557. Inyet another alternative embodiment, the external heat mechanism 559 maybe positioned in a recess or cavity within the build plate 557. Theexternal heat control mechanism 559 in another unshown example isarranged in a fixed positional relationship with respect to the buildplate 557. In yet another example, the external heat control mechanism559 may be positioned as close as possible to the formation of the newlayer on the build surface 512 to control the temperature profilethereof (as explained further herein). For example, the external heatcontrol mechanism 559 may be positioned as close as possible to thebottom surface 566 of the build plate 557. In another embodiment, theexternal heat control mechanism 559 may be positioned proximate to thebottom surface 566 of the build plate 557 and include soft magneticmaterial that is configured to concentrate flux toward the build surface512 to control the temperature profile of the layers formed thereon. Asshown in FIGS. 5D and 5E, the external heat control mechanism 559 mayform an interior space or void that is substantially aligned with theaperture 564 of the build plate 557 (e.g., in the vertical direction).The intermediate 516 may extend through the interior space or void ofthe external heat control mechanism 559 and into the aperture 564 of thebuild plate 557. Stated differently, a portion of the external heatcontrol mechanism 559 may extend at least partially about theintermediate 516. The component translating mechanism 560 may thereby beoperable to translate the intermediate 516 with respect to the externalheat control mechanism 559 (and the build plate 557).

The external heat control mechanism 559 may be operable to form apredetermined temperature profile of the build surface 512. For example,the external heat control mechanism 559 in one example includes at leastone induction coil that substantially surrounds the intermediate 516when the intermediate 516 is positioned within the aperture 564 of thebuild plate 557. As the airfoil 510 is conductive and so is the to-bemanufactured tip, the at least one induction coil of the external heatcontrol mechanism 559 is able to control the temperature of theintermediate 516 and the tip as electric current is passed through thecoil and a magnetic field is created. Further, as the external heatcontrol mechanism 559 is positioned proximate to the bottom surface 566of the build plate 557, the external heat control mechanism 559 iscapable of controlling the temperature of the build surface 512 toensure that the layers of the blade tip formed by the metallic powder“P” are not cracked. In this way, the external heat control mechanism559 is operable to form a predetermined temperature profile of the buildsurface 512 to prevent cracking of the blade tip.

In one example, a predetermined temperature profile is generated of atleast one newly, additively formed layer from the sintering or fusiontemperature of the at least one layer to the solidification temperaturethereof (e.g., about 1300° C., depending upon the composition of themetal alloy powder “P”) such that, at least upon solidification, the atleast one layer is crack-free. The predetermined temperature profile ofa newly formed layer, such as a predetermined cooling profile from thesintering or fusion temperature to the solidification temperaturethereof, that results in the solidified layer being crack free may beempirically determined, experimentally determined or a combinationthereof In some embodiments, the predetermined temperature profile maybe a range of predetermined cooling profiles of at least one newly andadditively formed layer from the sintering or fusion temperature to thesolidification temperature thereof such that, at least uponsolidification, the at least one layer is crack-free. A particulartemperature profile made from at least one newly and additively formedlayer that is effective in preventing cracks in the at least one layerat least upon solidification may be influenced or depend (at least inpart) by a number of factors, such the composition of the metallicpowder “P”, the thickness of the at least one layer, theshape/configuration of the at least one layer, the initial temperatureof the at least one layer (i.e., the fusion temperature), thesolidification temperature of the at least one layer, the temperaturegradient between the at least one layer and the preceding and/orsubsequent layer or formed portion, the desired microstructure of the atleast one layer after solidification, the ultimate operating parameterof the additively manufactured tip, the desired speed of the formationof the at least one layer (i.e., the movement of the component by thecomponent translating mechanism 560), etc. In one example, the apparatus550 forms or applies the predetermined temperature profile to an endportion of the airfoil 510 (e.g., to at least one newly formed layer),such as a cooling profile from fusion to solidification, by use of atleast the external heat control mechanism 559 and the componenttranslating mechanism 560 (to translate the intermediate 516 withrespect to the external heat control mechanism 559).

The predetermined temperature profile is typically greater than ambienttemperature and less than a temperature required to melt the depositedmaterial, i.e. 200-1200° C., preferably 500-1200° C., more preferably1000-1200° C. . In one embodiment, the external heat control mechanismsustains the temperature during the DMLM at about 1000° C. Heating atsuch a temperature promotes growth of the crystalline grains formed inthe blade tip, thereby allowing multiple ultra-thin additive layers(˜20-100 μm thick) to be formed that in turn, result in a blade tiphaving improved feature resolution. In addition, heating to suchtemperatures avoids cracking of the deposit, which occurs at lowertemperatures.

As shown in FIG. 5D, a method of manufacturing a blade tip with theapparatus 550 includes translating the blade tip, such as via thecomponent translating mechanism 560, with respect to the build plate 557such that the build surface 512 of the intermediate 516 positionedwithin the aperture 564 (potentially below the top surface 568 of thebuild plate 557) and in engagement with the seal 558. The aperture 564of the build plate 557, the seal 558 and the build surface 512 maycooperate to form the powder bed 565 for holding the metallic powder“P”. During such a condition, the powder supply mechanism 554 may exposemetallic powder “P”, as also shown in FIG. 5D. With metallic powder “P”exposed, the powder translating mechanism 555 may then fill the powderbed 565 by depositing the exposed metallic powder “P” through theaperture 564 and over the seal 558 and the build surface 512. The powderbed 565 may thereby form a layer of metallic powder “P” over or on thebuild surface 512. As noted above, in one example, the thickness of thelayer of metallic powder “P” over or on the end portion 64 of thecomponent “C” is within the range of 30-50 μm.

Once metallic powder “P” is deposited within the powder bed 565 and alayer of metallic powder “P” is thereby formed over or on the buildsurface 512, as shown in FIG. 5D the directed energy source 569 and thebeam directing mechanism 570 may direct a beam of energy to the layer ofdeposited metallic powder “P” in a pattern to fuse the metallic powder“P” to the build surface 512 as a new cross-sectional layer. After thenew cross-sectional layer is formed on the build surface 512, theexternal heat control mechanism 559 is used to form a temperatureprofile of at least the newly formed cross-sectional layer to preventcracking. Also after the new cross-sectional layer is formed on thebuild surface 512, and potentially during or part of the formation ofthe temperature profile of at least the newly formed cross-sectionallayer, the blade tip (not shown in FIG. 5D) may be translated withrespect to the build plate 557 and the external heat control mechanism559 by the component translating mechanism 560. The blade tip may betranslated to a lower position in the powder bed 565 such that the buildsurface 512 with the newly formed layer is positioned within theaperture 564 (potentially below the top surface 568) and in engagementwith the seal 558. The build surface 512 may then be in a condition fordeposition and fusion of metallic powder “P” in the powder bed 565 toform another layer thereon. In this way, translating the blade tip,depositing the metallic powder “P”, fusing the metal powder “P” layer onthe build surface 512 of the metal-ceramic intermediate 516, and formingthe temperature profile may form a cycle that may be performed aplurality of times to manufacture or form the metallic blade tip in across-sectional layer by cross-sectional layer fashion.

The beam directing mechanism 570 moves or scans the focal point of anunfocused laser or electron beam emitted by the directed energy source569 across the build surface 512 during the DMLM processes. The beamdirecting mechanism 570 in DMLM processes is typically of a fixedposition but the optical (e.g. telecentric lenses, mirrors, beamsplitters) or electronic (e.g. deflector coils, focusing coils)contained therein may be movable in order to allow various properties ofthe laser beam to be controlled and adjusted. However, in someembodiments the beam directing mechanism 570 itself may be moved todifferent positions for such adjustments. The speed at which the laseris scanned is a critical controllable process parameter, impacting howlong the laser power is applied to a particular spot. Typical laser scanspeeds are on the order of 10 s to 100 s of millimeters per second.

In certain embodiments the directed energy source 569 is a diode fiberlaser array (e.g. a diode laser bar or stack) that includes a pluralityof diode lasers or emitters that each emit a beam of radiation. Acylindrical lens may be positioned between the diode lasers and aplurality of optical fibers. The cylindrical lens compensates for thehigh angular divergence in the direction perpendicular to the diodejunction of the lasers, typically reducing the beam divergence in thefast axis to less than that of the slow axis, thereby easing theassembly tolerances of the overall system compared to an assembly whichdoes not use any coupling optics (i.e., one in which each fiber issimply placed in close proximity to the laser to which it is to becoupled). However, it should be appreciated that diode fiber laserarrays that do not use coupling optics may be used with the presenttechnology. In certain embodiments, the plurality of optical fibers mayfurther includes lenses at their respective ends that are configured toprovide collimated or divergent laser beams from the optical fibers. Itshould also be appreciated that even in the absence of these lenses, theends of the optical fibers 109 may be adapted to provide collimated ordivergent laser beams.

As an alternative to induction heating, the DMLM process of the presentinvention may be equipped with radiant heating, where at least a portionof the blade tip that is being additively manufactured and a retainingwall that defines a build chamber are heated to a desired temperature,such as the additive manufacturing processes disclosed in U.S. PatentApplication No. 2013/0101746 (currently assigned to Aeroj et Rocketdyne,Inc.), which is incorporated herein by reference in its entirety. Aplurality of heating elements may be mounted or supported upon theretaining wall. Preferably, at least the additively manufactured bladetip and the retaining wall proximal to the unfocused irradiation beamare radiantly heated. More preferably, the blade tip and the entirety ofthe workspace within the build chamber are subject to radiant heating.The heating elements generate a radiant heat that maintains the entireworkspace at a desired temperature. The desired temperature is typicallygreater than ambient temperature and less than a temperature required tomelt the deposited material.

The DMLM apparatuses and methods described above may be used toconstruct all or part of the blade tip, potentially in combination withother methods. For example, to construct all of blade tip via theapparatus 550 and methods described above, a seed component mayinitially be utilized for the formation of a first layer thereon. Inother embodiments, to construct part of the blade tip via the apparatus550 and methods described above the layers may be formed on apre-existing partially formed tip.

In one particular embodiment, a DMLM apparatus or system in accordancewith the present invention (with or without induction/radiant heating)may be combined with another additive manufacturing system andtechnique, namely the direct ink writing (DIW), which is also known as“robocasting”, for computer-controlled deposition of ceramic slurry. Asillustrated in FIG. 5F, a DIW robot or system 571 (representedschematically) extrudes or squeezes a filament of “ink”, which in thiscase is ceramic slurry, from a nozzle or syringe 572 to form a ceramicpart layer by layer (the extruded ceramic slurry eventually dries andsolidifies). The DIW process may be utilized, for example, to re-formthe ceramic core and/or ceramic shell, or to refine the build surface512 after the machining step. However, it should be noted that thepresent invention is not restricted to DIW and that any suitablecomputer-controlled ceramic additive manufacturing technique may be usedin combination with DMLM.

It is shown in FIGS. 5A-5D that the ceramic outer shell 502 remainsattached to the airfoil 510 after the casting process and even duringthe subsequent additive manufacturing of the blade tip. The shell may bealternatively removed after the casting is complete and prior to themanufacturing of the tip (core retained). However, retaining the shellmay provide additional advantages, such as providing improved visualcontrast to the step of removing a portion of the airfoil 510-ceramiccasting mold 500 combination to create the build surface 512. Visualcontrast improves automated recognition and registration of the bladetip. A robot can find each (slightly different) blade tip and adjust theDMLM program accordingly. The retained shell can also provide uniformheating or insulation of the parent airfoil 510 during a heated DMLMprocess, which promotes crystalline grain growth.

Prior to any of the additive manufacturing processes described herein,the ceramic-metal build surface may be treated for ceramiccontamination, e.g. by vacuuming, chemical treatment with an alkalinemedia, or with a protective layer covering the build surface or acombination of two or more of these, which promotes clean bonding of theadditive metal to the cast parent airfoil.

The ceramic cores of the present invention may incorporate specialdesign features, for example one or more witness features. In someembodiments, the witness feature is embedded in the ceramic core in sucha way that only one surface of the witness feature is shown without thefeature sticking out from the core surface. In other words, the witnessfeature is co-planar with the core surface. Alternatively, the witnessfeature forms a notch protruding from the core surface. In certainembodiments, the witness feature is made of a material other thanceramic. In other embodiments, the witness feature, like the core, isalso made of ceramic, which together with the core may be additivelymanufactured, e.g. using one or more of the techniques disclosed in theApplicant's co-pending application Ser. Nos. 15/377,673; 15/377,796;15/377,728; 15/377,759; 15/377,787; 15/377,746; 15/377,766; and15/377,783. The disclosures of each of these applications areincorporated here by reference in their entireties. Alternatively, thenotch witness feature and the ceramic casting mold are manufacturedusing the technique selective laser activation (SLA). For example, thenotch may be formed in the SLA metallic mold that is later used to formthe wax core.

FIG. 6A shows a ceramic casting mold 600 that includes an outer shell602 and an inner core 604 that define one or more inner cavities 606.The inner core 604 further includes a witness feature 614. After theinner cavities 606 are filled with a liquid metal 608 to form an airfoil610, a portion of the airfoil 610-ceramic casting mold 600 is removeduntil the witness feature 614 is exposed (see, FIG. 6B). Alternatively,a portion of the airfoil 610-ceramic casting mold 600 is removed untilthe witness feature 614 is exposed and also removed. Yet alternatively,a portion of the airfoil 610-ceramic casting mold 600 is removed quicklyand when the witness feature 614 is exposed, the removal process isslowed down and the witness feature 614 helps to determine how much moreof the airfoil 610-ceramic casting mold 600 needs to be removed. In thatway, having one or more witness features in the ceramic core enables arepeatable location of the first additive plane.

In another embodiment, ceramic cores of the present invention includefeatures that, upon removal of the core, correspond to a plurality ofturbulators being formed at the underside of the tip cap. These as-castturbulators promote and/or increase heat transfer between the heatedsidewalls of the airfoil and the internal cooling air. FIGS. 7A and 7Bpresent a cutaway view of a partial turbine blade 700 with the aft sideof the blade looking forward, showing the airfoil 702, airfoil outerwall 706, trailing edge 710, tip 704 and tip wall 708. The airfoil 702and the tip 704 are separated by the tip cap 712 that has a plurality ofturbulators 714 oriented circumferentially (FIG. 7A) or axially (FIG.7B).

In an alternative embodiment, the tip cap having a series ofheat-transferring turbulators as described above is additivelymanufactured rather than cast.

Yet other specific features or contouring may be incorporated in theceramic core design of the present invention, including but notnecessarily limited to integrated ceramic filaments between the core andshell of the mold that can be utilized to form holes, i.e., effusioncooling holes, in the cast component made from these molds. The use ofsufficient ceramic filaments between core and shell to both locate andprovide leaching pathways for the core serpentine also enables theelimination of ball braze chutes. Ceramic filaments between the tipplenum core and the shell may also be provided to support a floating tipplenum, eliminating the need for traditional tip pins, and theirsubsequent closure by brazing.

Conventional ceramic cores, like the core 800A shown in FIG. 8A, includea tip plenum 802 having tip pins 804 which upon leaching form tip holeswithin the cast airfoil 810A that must subsequently be brazed shut. Inthe present invention, ceramic cores with an open tip casting (i.e. tipplenum and tip pins eliminated) such as the core 800B of FIG. 8B may beprovided. The open tip design results in improved control of the airfoilwall (e.g. thickness), faster and easier leaching of the ceramic coreand eliminates the need for brazing. As shown in FIG. 8B, after moltenmetal is poured into the ceramic core 800B, cooled and solidified toform a cast airfoil 810B, a portion at the top of the ceramic core 800Bmay be removed to create a planar surface for subsequent DMLMmanufacturing of the blade tip.

The blade tips of the present invention may incorporate one or moreholes formed thereon during the additive manufacturing process. Theceramic core may be leached through these holes.

FIG. 9 depicts a method of repairing a field-return turbine blade with adamaged tip in accordance with the present invention. A field-returnairfoil 900 having an outer wall 902 and inner void 904 is provided inFIG. 9A. A ceramic slurry is injected into the void 904, which thenhardens to re-form the internal ceramic core 906. In certainembodiments, the external ceramic shell 908 enveloping the entire blade900 may also be re-formed. Subsequent steps in the repair are similar tothose described in the new-cast turbine blade manufacturing methods,i.e. removal of the top portion of the airfoil and the core to reveal abuild surface and additive manufacturing of a replacement blade tip.

In yet another aspect, the present invention provides an alternativemethod of repairing a field-return turbine blade with a damaged tip thatdoes not utilize a ceramic-metal build surface for the additivemanufacturing of a replacement tip. Referring to FIG. 10, an ultra-thinshim 1012 is placed on top of the damaged blade tip or the airfoil 1010at step S1002, which offers a flat or planar surface 1012 to support thesubsequent additive manufacturing of a new or replacement blade tip atstep S1004 (with optional induction or radiant heating as disclosedherein). In some embodiments, a ceramic core and/or a ceramic shell (notshown in FIG. 10) may be present prior to the additive manufacturing ofthe blade tip.

The ultra-thin shim 1012 is preferably made of the same material as themetallic powder used to additively manufacture the blade tip.Importantly, due to its thickness of merely less than 1000 μm, forexample 25-900 μm (0.025-0.9 mm), preferably 200-750 μm (0.02-0.75 mm)and more preferably 250-500 μm (0.25-0.5 mm), the ultra-thin shim 1012can be more readily and more easily joined to the outer wall of theairfoil 1010 compared to the thin plate 50 and the shim 50′ disclosed inthe earlier-referenced U.S. Patent Application Publication No.2010/0200189. The joining may be done using the same processes describedherein. Also importantly, the ultra-thin shim 1012 is not limited by itsgeometrical shape, but preferably it fully covers the outer walls of theairfoil 1010 to form a continuous and sealed support surface with one ormore overhangs 1016 for the additive manufacturing of the blade tip. Theoverhang(s) 1016 is subsequently removed at step S1006, for example, bymachining or grinding or any other known equivalent technique.

In an aspect, the present invention relates to the manufacturing methodsof the present invention incorporated or combined with features of othermanufacturing methods, apparatuses and ceramic core-shell molds. Thefollowing patent applications include disclosure of these variousaspects of these methods and molds:

U.S. patent application Ser. No. 15/406,467, titled “AdditiveManufacturing Using a Mobile Build Volume,” with attorney docket number037216.00059, and filed Jan. 13, 2017.

U.S. patent application Ser. No. 15/406,454, titled “AdditiveManufacturing Using a Mobile Scan Area,” with attorney docket number037216.00060, and filed Jan. 13, 2017.

U.S. patent application Ser. No. 15/406,444, titled “AdditiveManufacturing Using a Dynamically Grown Build Envelope,” with attorneydocket number 037216.00061, and filed Jan. 13, 2017.

U.S. patent application Ser. No. 15/406,461, titled “AdditiveManufacturing Using a Selective Recoater,” with attorney docket number037216.00062, and filed Jan. 13, 2017.

U.S. patent application Ser. No. 15/406,471, titled “Large ScaleAdditive Machine,” with attorney docket number 037216.00071, and filedJan. 13, 2017.

U.S. Patent Application No. [ ], titled “METHOD OF REPAIRING TURBINECOMPONENT” with attorney docket number 037216.00094/315780A, and filedFeb. 22, 2017;

U.S. Patent Application No. [ ], titled “METHOD OF MANUFACTURING TURBINEAIRFOIL WITH OPEN TIP CASTING AND TIP COMPONENT THEREOF” with attorneydocket number 037216.00091/315780B, and filed Feb. 22, 2017;

U.S. patent application Ser. No. [ ], titled “METHOD OF MANUFACTURINGTURBINE BLADE TIP” with attorney docket number 037216.00092/315780C, andfiled Feb. 22, 2017;

U.S. patent application Ser. No. [ ], titled “METHOD OF MANUFACTURINGTURBINE AIRFOIL AND TIP COMPONENT THEREOF USING CERAMIC CORE WITHWITNESS FEATURE” with attorney docket number 037216.00090/315780D, andfiled Feb. 22, 2017; and

U.S. patent application Ser. No. [ ], titled “METHOD OF REPAIRINGTURBINE COMPONENT USING ULTRA-THIN SHIM” with attorney docket number037216.00101/315780E, and filed Feb. 22, 2017.

The disclosures of each of these applications are incorporated herein intheir entireties to the extent they disclose additional aspects ofcore-shell molds and methods of manufacturing that can be used inconjunction with the core-shell molds disclosed herein.

This written description uses examples to disclose the invention,including the preferred embodiments, and also to enable any personskilled in the art to practice the invention, including making and usingany devices or systems and performing any incorporated methods. Thepatentable scope of the invention is defined by the claims, and mayinclude other examples that occur to those skilled in the art. Suchother examples are intended to be within the scope of the claims if theyhave structural elements that do not differ from the literal language ofthe claims, or if they include equivalent structural elements withinsubstantial differences from the literal language of the claims.Aspects from the various embodiments described, as well as other knownequivalents for each such aspect, can be mixed and matched by one ofordinary skill in the art to construct additional embodiments andtechniques in accordance with principles of this application.

1. A method of manufacturing a metal object, comprising: (a) pouring aliquid metal into a ceramic casting mold to form a cast component uponsolidification of the liquid metal, wherein the ceramic casting moldcomprises a ceramic core that comprises a witness feature and whereinthe ceramic core fills an inner cavity of the cast component; (b)removing a portion of the ceramic casting mold and the cast componentuntil the witness feature is revealed to create a surface portion ofcast component and a surface portion of the ceramic core; (c) depositinga layer of metallic powder onto the surface portions; (d) irradiating atleast a portion of the metallic powder to form a fused layer; and (e)repeating steps (c)-(d) until the metal object is formed.
 2. The methodaccording to claim 1, wherein the surface portions are planar surfaces.3. The method according to claim 1, wherein the ceramic casting moldfurther comprises a ceramic shell and at least one inner cavity betweenthe shell and the core.
 4. The method according to claim 3, furthercomprising removing the ceramic shell before step (b).
 5. The methodaccording to claim 1, further comprising treating the surface portionsto prevent ceramic contamination.
 6. The method according to claim 1,wherein steps (c)-(e) are carried out in the presence of inductionheating, radiant heating or a combination of both.
 7. The methodaccording to claim 1, further comprising removing the ceramic castingmold from the metal object.
 8. The method according to claim 1, whereinthe metal object is a turbine blade or vane, and the cast component isan airfoil.
 9. A method of manufacturing a turbine blade or vane,comprising: (a) pouring a liquid metal into a ceramic casting mold toform a cast component upon solidification of the liquid metal, whereinthe ceramic casting mold comprises a ceramic core that comprises awitness feature and wherein the ceramic core fills an inner cavity ofthe cast component; (b) removing a portion of the ceramic casting moldand the cast component until the witness feature is revealed to create asurface portion of cast component and a surface portion of the ceramiccore; (c) depositing a layer of metallic powder onto the surfaceportions; (d) irradiating at least a portion of the metallic powder toform a fused layer; and (e) repeating steps (c)-(d) until the blade orvane is formed.
 10. The method according to claim 9, wherein the surfaceportions are planar surfaces.
 11. The method according to claim 9,wherein the ceramic casting mold comprises an external shell and atleast one inner cavity between the shell and the core.
 12. The methodaccording to claim 11, further comprising removing the external shellbefore step (b).
 13. The method according to claim 9, further comprisingtreating the surface portions to prevent ceramic contamination.
 14. Themethod according to claim 9, wherein steps (c)-(e) are carried out inthe presence of induction heating, radiant heating or a combination ofboth.
 15. The method according to claim 9, further comprising removingthe ceramic casting mold from the blade or vane.
 16. The methodaccording to claim 9, wherein the cast component is an airfoil.
 17. Acast metal turbine blade or vane workpiece, comprising: a metallic bladeor vane portion with a ceramic core that comprises a witness feature;and a substantially planar top surface of the workpiece that comprises ametallic portion and a ceramic portion.
 18. The cast metal turbine bladeor vane workpiece according to claim 17, wherein the ceramic core fillsan inner cavity of the metallic blade or vane.
 19. The cast metalturbine blade or vane according to claim 17, wherein the blade or vaneis further surrounded by a ceramic shell.
 20. The cast metal turbineblade or vane according to claim 17, further comprising a protectivelayer for preventing ceramic contamination.